Thermal control systems have been devised to regulate the temperature of waste heat generating equipment. These systems typically dissipate heat from the equipment by mounting the equipment conductively against an external surface such as the exterior wall of a container, spacecraft or airplane and, thereby, allowing the external surface to radiate heat into the surrounding environment. Other designs for control systems allow the equipment to radiate waste heat into the interior of the vehicle and then use an external radiating surface such as the outer walls or discrete radiators mounted on the vehicle's exterior to cool the interior. In other systems thermal switches or heat pipes conduct waste heat from the interior of the vehicle to discrete, external radiators which discharge the waste heat directly into the surrounding environment. Heat transfer through the heat pipes and radiators has, in some systems, been made variable by means of variable conductance heat pipes, thermal switches or thermocouple actuators for positioning louvers to cover the radiating surfaces.
With the exception of those designs having the heat generating equipment mounted directly against an external radiating surface, previously available thermal control systems have had relatively low thermal efficiencies due to the intrinsically low transfer efficiencies of non-conductive intermediate links between the equipment and the radiating surfaces. These systems have performed satisfactorily despite their thermal inefficiencies because the heat generating equipment has usually been operated at temperatures above 30.degree. C. over wide ranges of tolerance and the quantity of heat generated has been relatively small in comparison to the areas of the spacecraft available for use as external radiating surfaces. Their thermal inefficiencies, however, makes these systems unsuitable for reliably controlling the temperature of equipment operated below 30.degree. C. Additionally, they are unsuitable for maintaining precise equipment temperature in the presence of such external causes of thermal variation as earth shine, albedo and diurnal ambience. Furthermore, these systems depend upon large radiating surface areas and usually require deployable radiator panels that render them unsuitable for compact spacecraft applications.
Several features of spacecraft, particularly their low-temperature, gravity-free environment and their accessibility to a constant, albeit limited, source of energy in the form of solar radiation have prompted proposals for using earth-orbiting spacecraft as plants for carrying manufacturing equipment and processes such as a continuous flow electrophoresis system for manufacturing blood plasma. These proposals include deployment of spacecraft containing process plants via a Shuttle-type cargo vehicle for multi-year operations with periodic, interim visits to the orbiting spacecraft to remove finished products and deliver additional raw materials. Economy of scale requires, however, that to achieve economic viability, such plants must be much larger than previously deployed experimental and communications spacecraft. Spacecraft containing exothermic industrial processes will, accordingly, require thermal control systems having a capacity approximately proportional to the cubic power of the waste heat removing capacity of current spacecraft systems because the quantity of waste heat generated is proportional to the volume of the plane. Such systems must be able to continuously radiate between several hundred to a few thousand watts of waste heat. The multi-layer lifetimes planned for these plants and the current lack of cooling system machinery such as pumps and compressors, able to provide continuous, maintenance-free operation over such lifetimes necessitates a passive system having no moving parts while conservation of the available power obtained from solar energy for more profitable use in the industrial process requires that the system consume little or no power. The size of the cargo bay capacity in Shuttle-type vehicles limits the overall size of industrial process spacecraft, thereby providing an incentive for maximizing the volume available to the industrial process by miniziming that required by auxiliary plant components such as the thermal control system. Ease of handling during deployment and interim vists, a need to provide a dynamically stable spacecraft to avoid interruption of an industrial process with sudden G-forces, and maintenance of an alignment between the spacecraft's solar panels and the sun, further requires that the thermal control system be relatively compact and have its mass symmetrically distributed so as to not significantly effect stability of the spacecraft. Moreover, the presence of solar panels means that the heat rejection system cannot use deployable radiator panels that might shade the solar panels and thereby interrupt the flow of solar energy to the process.
There has not been yet developed a controllable system without moving parts for removing variable quantities of heat from a continuously flowing, closed process loop. Such a system would be very useful in an airborne or an outer space environment where corrective maintenance is not available. The inefficiencies in the thermal conduction provided and the relatively large exterior radiating surfaces required by currently available thermal control system designs means that these designs are unsuitable for use with proposed air or spaceborne processes. Moreover, the construction of these thermal control systems as integral parts of a spacecraft renders them suitable for use in space as a compact, quickly detachable part of a spacecraft. Consequently, the large size, periodic maintenance requirements, difficulties of disassembly, and thermal conduction inefficiencies of state of the art thermal control systems means that both a process plant and its thermal control system must be retrieved from space together, a requirement which unnecessarily restricts the size of the plant and, thereby, limits its profitability.